1. Field of Endeavor
The present invention relates to gas turbines, and more specifically a blade for a gas turbine.
Especially, the invention relates to designing rotor blades of an axial-flow turbine used in a gas turbine unit. The turbine rotor includes a rotating shaft with axial fir-tree type slots where several blade rows and several rotor heat shields are installed alternately one after another.
2. Brief Description of the Related Art
The schematized section of a gas turbine stage is shown in FIG. 1. The turbine 10 of FIG. 1 includes a stator 12 and a rotor 11. The stator 12 represents a housing and includes a vane carrier 15 with stator heat shields S1-S3 and vanes V1-V3 mounted therein. The stator 12 concentrically surrounds the rotor 11 and defines a hot gas path 13. Hot gas 16 generated in a combustion chamber (not shown) passes through profiled channels between the vanes V1-V3, hits against blades B1-B3 mounted in shaft slots of a rotor shaft 14, and thus makes the turbine rotor 11 rotate.
Inner platforms 23 of the 1st, 2nd and 3rd stage blades B1, B2 and B3, in combination with intermediate rotor heat shields R1, R2, form the inner outline of the turbine flow or hot gas path 13, which separates the cavity of rotor cooling air transit (cooling air 17) from the hot gas flow 16. To improve tightness of the cooling air flow path between adjacent blades in the circumferential direction, sealing plates 29 are installed. When cooling the rotor shaft 14, cooling air 17 in this design flows in the axial direction along a common flow path between blade roots 24 and rotor heat shields R1, R2 and enters in turn into the internal cavity (cooling channels) of the blade B1, then into that of the blade B2 and that of blade B3 (cooling air 18).
Turbine blades used in present day efficient gas turbine units are operated under high temperatures with minimum possible air supply. Striving towards cooling air saving results in complication of internal blade channel configurations. Therefore, the blade manufacturing process is very complicated. After blade casting, a problem frequently occurs in elimination (etching out) of a ceramic (casting) core from the blade internal cavity (cooling channels).
FIGS. 2 and 3 show the external configuration and internal channel geometry, respectively, of a typical gas turbine blade according to the state of the art. The blade 19 includes an airfoil 20 with a leading edge 21 and trailing edge 22, and a blade root 24 with an inlet 25 for supplying the internal cooling channel structure (FIG. 3) with cooling air. Blade root 24 and airfoil 20 are separated by a platform 23. The internal cooling channel structure includes a plurality of cooling channels 20 and 27a-c, which extend in the longitudinal direction of the blade 19. Usually, some parallel cooling channels 27a-c are connected in series to build one meandering channel, as is shown in FIG. 3. Such a meandering channel 27a-c results in a blind tube or dead end zone 28, which rules out any possibility that a liquid flow-through could be established to remove (by wet etching) ceramic core rests from there; this fact makes the manufacturing process more expensive and sets up a danger concerning the presence of detrimental remains of the core in internal blade channels.
If the blade cooling scheme of the gas turbine blade in question cannot be simplified without generating significant cooling air losses, then a technological possibility for a guaranteed and complete removal of the ceramic core from the internal blade cavity should be provided.